The disclosed embodiments generally pertain to a gas turbine engine. More particularly, but not by way of limitation, present embodiments relate to aircraft engine architecture having a reverse rotation integral drive and a vaneless turbine with counter-rotating blades.
A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber, a turbine, and a nozzle at the aft end of the engine. It will be readily apparent from those skilled in the art that additional components may also be included in the engine, such as, for example, low pressure and high pressure compressors, and high pressure and low pressure turbines. This, however, is not an exhaustive list. An engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.
In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy.
Various components of the gas turbine engine operate at highest efficiencies at different speeds. Many present aircraft engines utilize systems which directly couple the turbine and compressor and fan on shafts so that while one of the core components may be operating at maximum efficiency, the other components coupled by the same shaft are not operating at best efficiency. It is highly desirable to operate the aircraft engines at an efficient high bypass ratio, low fan pressure ratio with optimal design speeds for the fan, compressor or booster and the low pressure turbine.
It would also be desirable to reduce or minimize the number of stages required, by the low pressure compressor or booster, in the low pressure turbine.
As may be seen by the foregoing, it would be desirable to overcome these and other deficiencies with gas turbine engines components.